FIG. 1 shows an example of a combustion turbine engine 10. The turbine engine 10 includes a compressor section 12, a combustor section 14, and a turbine section 16. The combustor section 14 can include a plurality of combustors 15 (only one of which is shown) arranged in an annular array around a rotor 17. The turbine section 16 includes alternating rows of stationary airfoils 18 and rotating airfoils 20.
In operation, air is drawn in through the compressor section 12, where it is compressed and driven towards the combustor section 14. The compressed air 21 can be directed to the combustor 15 through an air intake 22. The air 21 can then be mixed with fuel to form an air/fuel mixture. In the combustor 15, the fuel/air mixture can be ignited to form a working gas. A duct 26 (sometimes referred to as a transition) can be provided for each combustor 15 to route the working gas to the turbine section 16.
Each transition duct 26 includes a transition body 28, an outer peripheral surface 33, an inlet end 30 and an exit end 32. The duct 26 further includes an outlet region 31 that includes the exit end 32. The transition duct 26 can be supported at various locations along its length. For example, the exit end 32 of the transition 26 can be supported to counter its own weight as well as to counter some of the forces imposed by the combustion gases flowing through the transition 26. In addition, the transition exit end 32 can be supported to allow proper alignment with the first row of stationary airfoils 18 in the turbine section 16.
One known support system 35 for the exit end 32 of each transition 26 is shown in FIG. 2. The system 35 is integral with each transition 26 and includes a support bracket 36 with seals 34. The support bracket 36 is provided at the exit end 32 of the transition 26 and is used to attach the transition 26 to a stationary structure 40 in the turbine section 16, such as a first stage blade ring or vane carrier 42 (see FIG. 1). The seals 34 can engage the stationary structure 40 in the turbine section 16. It should be noted that, when installed, the exit frame ring 37 associated with one transition 26 abuts the exit frame ring 37 associated with a neighboring transition 26, thereby forming an exit frame ring interface 38 between the two transition ducts 26.
Through experience, such a support system 35 has proven to have a number of drawbacks, particularly with respect to sealing. For instance, because the interface 38 of adjacent exit frame rings 37 lies between transitions 26, there is potential for compressed air from the combustor section to penetrate the interface 38 and enter the gas path of the turbine section 16, which can adversely affect engine performance and emissions. Similar sealing concerns occur at the interface between the exit end 32 of the transition 26 and the first stage of the stationary structure 40. Leakages at these interfaces are difficult to predict due to large tolerances and large transient deflections; thus, complicated sealing issues are introduced into the design of the support system 35. Very precise tolerances are required to set the exit end 32 of each transition 26 in place so that the seals 34 can properly engage the stationary structure 40 of the turbine section 16.
The known system 35 also includes relatively complex and wear prone subcomponents, requiring more expensive manufacturing and repair techniques and raising concerns of system downtime. Moreover, the known support system 35 suffers from large thermal stresses at support attachment locations, uses expensive components, and involves difficult and labor intensive assembly.
Thus, there is a need for a transition exit support system that can minimize the above concerns.